Orbit space transportation and recovery system

ABSTRACT

A preferred In Orbit Transportation &amp; Recovery System (IOSTAR™) ( 10 ) includes a space tug powered by a nuclear reactor ( 19 ). The IOSTAR™ includes a collapsible boom ( 11 ) connected at one end to a propellant tank ( 13 ) which stores fuel for an electric propulsion system ( 12 ). This end of the boom ( 11 ) is equipped with docking hardware ( 14 ) that is able to grasp and hold a satellite ( 15 ) and as a means to refill the tank ( 13 ). Radiator panels ( 16 ) mounted on the boom ( 11 ) dissipate heat from the reactor ( 19 ). A radiation shield ( 20 ) is situated next to the reactor ( 19 ) to protect the satellite payload ( 15 ) at the far end of the boom ( 11 ). The IOSTAR™ ( 10 ) will be capable of accomplishing rendezvous and docking maneuvers which will enable it to move spacecraft between a low Earth parking orbit and positions in higher orbits or to other locations in our Solar System.

CROSS-REFERENCE TO RELATED APPLICATIONS

The Present Patent Application is a Divisional Application, and is basedon U.S. patent application Ser. No. 10/779,869, which was filed on 17Feb. 2004, which was issued as U.S. Pat. No. 7,216,834 on 15 May 2007.

U.S. patent application Ser. No. 10/779,869 is a Continuation-in-Part ofU.S. patent application Ser. No. 10/755,200, which was filed on 9 Jan2004, which was issued as U.S. Pat. No. 7,070,151 on 4 July 2006.

U.S. patent application Ser. No. 10/755,200 is a Continuation-in-Part ofU.S. application Ser. No. 10/736,887, which was filed on 15 Dec. 2003,which was issued as U.S. Pat. No. 7,216,833 on 15 May 2007.

U.S application Ser. No. 10/736.887 is a Continuation-in-Part of U.S.patent application Ser. No. 10/298/138, which was filed on 15 Nov. 2002,which is now abandoned.

U.S. application Ser. No. 10/298,138 is a Continuation-in-Part of U.S.patent application Ser. No. 09/918,705, which was filed on 30 Jul. 2001,which is now abandoned.

The Applicants claims the benefit of priority under Sections 119, 120and 363 of Title 35 of the United States Code for any subject matterwhich is commonly disclosed in the Present Patent Application and in

-   -   Pending U.S. patent application IOS9601 CIPD, Ser. No.        10/779,869, filed on 17 Feb. 2004;    -   Pending U.S. patent application IOS9601 CIPC, Ser. No.        10/755,200, filed on 9 Jan. 2004;    -   U.S. patent application IOS9601 CIPB, Ser. No. 10/736,887, filed        on 15 Dec. 2003, which has been issued as U.S. Pat. No.        7,070,151;    -   PCT International Patent Application IOS9601-B&C-PCT,        PCT/GB04/000378, filed on 29 Jan. 2004, now expired;    -   PCT International Patent Application IOS9601-CIPB-PCT,        PCT/US03/32748, filed on 10 Nov. 2003, now expired;    -   U.S. Patent Application ITS9601 CIPA U.S. Ser. No. 10/298,138,        filed on 15 Nov. 2002, now abandoned; and    -   U.S. Patent Application ITS9601, U.S. Ser. No. 09/918,705, filed        on 30 Jul. 2001, now abandoned.

FIELD OF THE INVENTION

The present invention relates to the field of spacecraft and satellites.More particularly, this invention provides a transportation and rescuesystem for moving objects in space between low Earth orbits, higherorbits and beyond.

FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT

None.

BACKGROUND OF THE INVENTION

Hundreds of man-made satellites are currently in orbit around the Earth.Over the next decade, governments and companies around the globe plan tolaunch hundreds of new spacecraft for a variety of communications,defense and remote sensing projects. The placement of satellites intoEarth orbit can cost many millions of dollars. A conventional launchinvolves a large multi-stage, single-use rocket to lift a satellite intoa geosynchronous orbit.

A general description of conventional nuclear-propulsion systems may befound in a text entitled A Critical Review of Space Nuclear Power andPropulsion, edited by Mohamed S. El-Genk, which was published by theAmerican Institute of Physics in 1994.

The U.S. Departments of Energy and Defense and NASA developed plans fora Generic Flight System for space-based defense systems and NASAexploration missions called SP-100 in the mid-1980's. The SP-100 wasdesigned to supply nuclear-power for military and civilian spacesystems. This early system was designed as a single-use power stage fora single, permanently attached payload; and was never configured for anyon-orbit rendezvous, docking or servicing missions. The SP-100 isdescribed in the SP-100 Technical Summary Report, which was prepared forthe U.S. Department of Energy by the Jet Propulsion Laboratory and theCalifornia Institute of Technology in September, 1994.

Various nuclear electric propulsion systems are described in apublication entitled Nuclear Electric Propulsion, A Summary of ConceptsSubmitted to the NASA/DoE/DoD Nuclear Electric Propulsion Workshop,which was held in Pasadena, Calif. on 19-22 Jun. 1990.

The Aerospace Division of the Olin Corporation proposed a small enginefor the small satellite community called the Small Upper Stage (SUS).The SUS was designed to accomplish low Earth orbit transfers, orbitcircularizations and plane changes using hydrazine propulsion.

TRW has patented several methods and apparatus intended for the spacetransportation market. In U.S. Pat. No. 4,471,926, Steel describes aTransfer Vehicle for Use in Conjunction with a Reusable Space Shuttle.This spacecraft has a propulsion system that uses a low-thrustbi-propellant liquid rocket engine to provide a soft, low-accelerationascent. In U.S. Pat. No. 4,575,029, Harwood and Love disclose aspacecraft for transporting a payload from a space shuttle in a lowaltitude parking orbit to an operational orbit. In U.S. Pat. No.4,943,014, Harwood and Love reveal their “soft ride” method for changingthe altitude or position of a spacecraft in orbit using a liquidbi-propellant engine.

In U.S. Pat. No. 4,664,344, Harwell describes an apparatus and method ofcapturing an orbiting spacecraft. This device comprises a relativelysmall mechanical probe and fixture operated by an astronaut during aspacewalk.

In an article entitled Topaz Two Proves to Be a Gem for InternationalTech Transfer, contained in Technical Applications Report from BallisticMissile Defense Organization, 1995, thermoionic reactors for space-basedpower generation are disclosed.

Prospects for Nuclear Electric Propulsion Using Closed-CycleMagnetohydrodynamic Energy Conversion, by R. Litchford et al. waspresented at the 12th Annual Advanced Space Propulsion Workshop inHuntsville, Ala. on 3-5 Apr. 2001.

J. Collins et al. disclose a Small Orbit Transfer Vehicle for On-OrbitServicing and Resupply which was presented at the 15th Annual Utah StateUniversity Conference on Small Satellites at Logan, Utah, 13-16 Aug.2001.

In U.S. Pat. No. 4,754,601, Minovitch discloses “a propulsion system forreusable space-based vehicles is presented wherein the propulsiveworking fluid is atmospheric gas.”

In U.S. Pat. No. 5,260,639, De Young et al. describe “a method ofsupplying power to a device such as a lunar rover located on a planetarysurface.”

In U.S. Pat. No. 6,213,700, Koppel discloses a “method [which] serves toplace a space vehicle, such as a satellite, on a target orbit such asthe orbit adapted to normal operation of the space vehicle and startingfrom an elliptical initial orbit that is significantly different from,and in particular more eccentric than the target orbit.”

In U.S. Pat. No. 6,357,700, Provitola describes “an spacecraft/airship,which uses buoyancy and thrusters to ascend into space with lifting gasas propellant or fuel for thrusters, which may be conventional thrustersor electric turbojets or ion thrusters.”

In U.S. Pat. No. 5,260,639, Basuthakur et al reveal “a satelliteassembly [that] is formed from any number of bus modules which have asubstantially common shape and interior space volume.”

In U.S. Pat. No. 6,478,257, Oh et al. describe “systems and methods thatemploy a phase change material to provide thermal control of electricpropulsion devices.”

In U.S. Pat. No. 3,825,211, Minovitch presents a “space vehicle [which]carries a vaporizable propellant . . . [E]nergy is transmitted to thevehicle while in space by a laser beam originating on the ground or someother body or satellite.”

In U.S. Pat. No. 6,364,252, Anderman discloses a method of using dwelltimes in intermediate orbits to optimize orbital transfers, as well asan apparatus for satellite repair.

In U.S. Pat. No. 6,669,148, Anderman et al. describes a method andapparatus for supplying orbital space platforms.

In U.S. Pat. No. 5,294,079, Draznin et al. describes a space transfervehicle.

The development of an in-orbit space transportation and rescue vehiclewould dramatically reduce the cost of changing the orbital position of asatellite. Such a system would revolutionize the military and commercialspace industries, and fill a long-felt need in the telecommunications,direct-broadcast and remote-sensing industries.

SUMMARY OF THE INVENTION

The In Orbit Space Transportation & Recovery System (IOSTAR™) willrevolutionize the commercial space industry by providing a lower costalternative to conventional methods of moving spacecraft in orbit.Instead of using a multi-stage rocket powered by expensive and dangerouschemical fuels to lift a payload to a geosynchronous or geostationaryorbit, the IOSTAR™ will rendezvous with a satellite waiting in a lowEarth orbit, dock with the satellite and then gently transport it to analtitude of 22,300 miles using reliable nuclear-powered electricpropulsion. The IOSTAR™ will also be available to relocate, rescueand/or retrieve satellites in need of repositioning or repair, and willbe capable of ferrying objects to the Moon and to the neighboringplanets of our Solar System.

One embodiment of the IOSTAR™ includes a collapsible boom which maydouble as a radiating surface, and which expands to its fully extendedposition after reaching orbit. The boom is connected at one end to atank which stores xenon which fuels ion propulsion engines located atthe opposite end of the boom. Docking hardware which is capable ofengaging a wide variety of objects in space is coupled to the farthestend of the boom near the fuel tank. A nuclear reactor, a radiationshield, an energy converter and a large array of heat-dissipatingflat-panel radiators are mounted on the boom between the reactor and apayload grasping device.

An appreciation of the other alms and objectives of the presentinvention and a more complete and comprehensive understanding of thisinvention may be obtained by studying the following description of apreferred embodiment and by referring to the accompanying drawings.

A BRIEF DESCRIPTION OF THE DRAWINGS

FIGS. 1A & 1B present top and end views of one of the preferredembodiments of the In Orbit Space Transportation & Recovery System(IOSTAR™) vehicle in its fully deployed, orbital configuration.

FIG. 2 depicts a separate service and refueling vehicle.

FIG. 3 is a side view of the present invention in its fully deployedconfiguration.

FIG. 4 reveals the present invention in a folded and collapsedconfiguration that may be loaded aboard a launch vehicle.

FIGS. 5, 6, 7 and 8 present side and end views of preferred embodimentsof the present invention stowed aboard a launch vehicle.

FIG. 9 is a block diagram of control systems installed in the IOSTAR™spacecraft.

FIG. 10 is a cross-sectional view of an ion propulsion engine utilizedby one embodiment of the IOSTAR™ spacecraft.

FIG. 11 is a cross-sectional view of a portion of one embodiment of theinvention inside a launch vehicle shroud.

FIG. 12 presents a diagram which provides an overview of the BraytonSystem, which is used as the energy converter in one embodiment of theinvention.

FIG. 13 supplies a perspective view of an alternative embodiment of theIOSTAR™.

FIG. 14 is a schematic depiction of the process of conveying a satellitefrom a low Earth orbit to a higher orbit using the present invention.

FIG. 15 illustrates a method for repositioning a satellite.

FIGS. 16 and 17 are comparisons of high orbit architectures forconventional and IOSTAR™ missions.

FIGS. 18, 19, 20 and 21 exhibit four IOSTAR™ missions.

FIGS. 22 and 23 show the IOSTAR™ and the Intentional Space Station.

FIG. 24 depicts a satellite having an array of antennas that could becombined with the IOSTAR to provide direct broadcast services.

FIG. 25 exhibits methods of supplying controlled kinetic energy, powerand information and data to satellites and celestial bodies.

A DETAILED DESCRIPTION OF PREFERRED & ALTERNATIVE EMBODIMENTS

I. Overview of Embodiments of IOSTAR™

FIGS. 1A and 1B reveal side and end views of one of the preferredembodiments of the In Orbit Transportation & Recovery System, or IOSTAR™10. IOSTAR™ is a Trade and Service Mark owned by the Assignee. IOSTAR™is a reusable spacecraft 10 which is designed primarily for orbitaltransportation and rescue services.

In this Specification and in the Claims that follow, the term“satellite” refers to any object in any orbit around any body, whethernatural or man-made. Any object which remains above the Earth'satmosphere for an extended time is a satellite, since it must be withinthe gravitational influence of the Earth or some other celestial body. A“celestial body” is any planet, moon, asteroid, comet, star, galaxy, orany other aggregation of matter. The terms “vehicle” and “spacecraft”concerns any man-made device or means used temporarily below orbitalaltitude, in orbit, or beyond the Earth's atmosphere, or for travel inspace; including a ship, structure, platform, machine or manufacturethat may travel beyond Earth's orbit. The term “orbit” generally means apathway or line of movement of an object that includes any position atany point or altitude above the surface of the Earth or other celestialbody which allows an object, satellite or spacecraft to move above theEarth's surface with or without aerodynamic lift, up to a distance whichis still within the Earth's gravitational field. Orbits may includepathways around the Earth, the Moon, the Sun or any other celestialbody.

In general, the term “low Earth orbit” encompasses any orbital altitudebelow geosynchronous or geostationary orbit. In general, the term “highEarth orbit” encompasses any orbital altitude from geosynchronous orgeostationary orbit to any position above geosynchronous orgeostationary orbit within the Earth's gravitational field. In general,the term “space” refers to any position generally outside the Earth'satmosphere. The term “object” pertains to any configuration, embodimentor manifestation physical mass or matter, including natural objects suchas asteroids or MMOD's (micro-meteoroids and orbital debris), man-madedevices, or other things or items. The term “payload” encompasses anyitem or cargo that is carried or transported. A typical payload is asatellite, but a payload could be a load of material, supplies,equipment, or some other object. A payload could also include a humancrew, and/or other living beings including plants and animals.

In one embodiment of the invention, the backbone or central skeleton ofthe IOSTAR™ 10 comprises a lightweight but strong, generally metallic orcomposite, collapsible, compressible or at least partially foldable boom11. The boom 11 provides structural support, but is also capable offitting inside a launch vehicle when collapsed, and then extending toits fully deployed length after launch. The launch vehicle may be asingle use vehicle, or may be reusable or expendable. In a preferredembodiment of the invention, the IOSTAR™ will be lifted into orbit bythe United States Space Shuttle.

In one embodiment, one end of the boom 11 is connected to an electricpropulsion system 12. In general, an electric propulsion system is anymeans which employs electromagnetic forces to generate thrust. In oneembodiment, a tank 13 which stores propellant for the electricpropulsion system 12 is connected to the boom 11 at the end oppositefrom the ion engines 12. In a preferred embodiment of the invention, theelectric propulsion system is an ion propulsion system 12 which expelsions to produce thrust. Table One contains a list of some of the varioustypes of electric propulsion systems that may be utilized to implementthe present invention.

TABLE ONE Electric Propulsion Alternatives. Electrothermal ArcjetsResistojets Electrothermal thruster Continuous wave Laser & LaserAblative Microwave heated thruster Electromagnetic Magnetoplasmadynamicthruster  Self-Field  Applied Field Hall effect thruster Stationaryplasma thruster θ-pinch thruster Compact toroid thrusterPulsed-inductive thruster Coil-gun Z-pinch discharge thruster Coax gunPulsed-plasma thruster Rail-gun Mass-driverElectrostatic Ion engineField emission Other Magnetic loop sail Electrodynamic TetherThe recitation of electric propulsion alternatives in Table One is notintended to exclude any unlisted or equivalent alternatives.

In a preferred embodiment of the invention, the ion engines 12 employxenon ions, so the tank is filled with xenon. In an alternativeembodiment, the ion propulsion system 12 includes a Hall thruster. Otherembodiments of the invention may employ different fuels, and may utilizemultiple fuels. The invention may utilize any tank means which holds,envelopes or stores suitable propellants.

In a preferred embodiment of the invention, the tank 13 is refillable,and may be refilled in a relatively low or zero gravity environment. Oneembodiment of the invention includes one or more tanks that provides thepropulsion system with propellant. In one embodiment of the invention,the tank may be refilled by a separate, automatic, unmanned spacecraftas shown in FIG. 2. When the IOSTAR™ vehicle runs low on propellant, itwill be replenished by a servicing vehicle that either transfers all itspropellant and is then released; or transfers its propellant graduallyand is released when empty. In one embodiment, the IOSTAR™ will have alower pressure tank so that pumping is kept to a minimum or eliminated.In another alternative, the size of the lower pressure tank is smaller,and includes limited life thrusters attached to the servicing vehicle.The electric thrusters on the service vehicle can be operated at higherpower than the rest of the thrusters on the IOSTAR™ to enhanceperformance since the high power reduces lifetime, the thrusters arereplaced with the next service vehicle. The thrusters may have a limitedlifetime, and be used for a relatively small number of missions, or, maylast for the entire lifetime of the IOSTAR™. The service vehicle may beequipped with application specific thrusters that are replaced with thenext service vehicle.

Table Two contains a list of some of the propellants that may beemployed to practice the present invention.

TABLE TWO Propellants Xenon Mercury Aluminum Bismuth Krypton HeliumArgon Production Kr—Xe mix Hydrogen Nitrogen N2 + 2H2 NH3 H2O NH3 CO2N2H4 CH4 Air Lithium Cesium Indium TeflonThe recitation of propellant alternatives in Table Two is not intendedto exclude any unlisted or equivalent alternatives.

The end of the boom 11 which holds the propellant tank 13 is equippedwith reusable docking hardware 14 that is able to contact or grasp asatellite 15 or some other object in space. In one embodiment of theinvention, the tank is replenishable. This docking hardware 14 may bereferred to as a grasping device, and may comprise any multiple-usemeans for engaging an object above the Earth. Many different embodimentsof the docking hardware 14 may be incorporated in the present invention.In general, the preferred embodiment of the invention is reusable,utilizes a multiple-use docking device 14, which, unlike some of theprior art, is designed for many missions over a relatively longlife-time.

The docking hardware 14 may be configured to interact with a widevariety of satellites 15 or other objects above the surface of theEarth. In general, the docking hardware 14 comprises any reusable ormultiple-use means which is adapted to interact with or to engage apayload launch vehicle interface, or to otherwise engage an object inspace. In one embodiment, which is shown best in FIGS. 1A, 3, 4, 5, 7and 13, the grasping means comprises a plurality of segments, whichextend outwardly away from one end of the boom. These segments areconfigured to partially surround or to enclose an object, and then toengage and to grasp a wide variety of satellites 15 or other objectsabove the surface of the Earth without the need for any preconfigureddocking interface on the object which is grasped. Unlike some previousequipment designed for launch into orbit, the present invention includesa grasping means 14 which is not permanently affixed or connected to apayload.

A radiator 16 is disposed generally perpendicular to the boom 11 nearthe ion thrusters 12. The radiator 16, which conveys a coolant throughmanifold 17 and fluid flow tubes 18, dissipates heat from an energyconverter out to space. The energy converter is powered by a nuclearreactor 19. The radiator 16 is generally situated between the graspingdevice 14 and the reactor 19. In general, the radiator 16 is a pumpedfluid loop. An alternative embodiment comprises a capillary pump loopand/or heat pipes. In another alternative embodiment of the invention,the radiator 16 may be disposed along the boom 11, or a single combinedradiator/boom means may be employed.

The reactor 19 generates heat through the controlled fission of nuclearfuel. This heat is then converted to electrical power. In a preferredembodiment, the reactor 19 is gas-cooled. In alternative embodiments,the reactor 19 employs a liquid-metal coolant, or some other workingfluid or heat pipes. The reactor 19 is coupled to a radiation shield 20,which protects the object, payload or satellite 15 from radiationgenerated by the reactor 19. In one embodiment of the invention, theradiator 16 is configured to remain entirely within the protective zoneof the radiation shield 20. In one embodiment, the radiation shield 20incorporates multiple zone shielding to minimize mass. In anotherembodiment, the radiation shield includes a recuperator that is alsoemployed as a gamma shield. The transport of a living payload, such aslive humans, may require additional levels of shielding.

In a preferred embodiment of the invention, from 250 kW to 500 kW ofsustained electrical power may be generated aboard an IOSTAR™, whichvastly exceeds the sustained power generating capabilities of any priorman-made satellite or spacecraft. This power generation capacity is hugewhen compared to the power levels of conventional satellites andspacecraft, which typically operate with less than 20 kW of power. Thisimmense on-orbit power generating capacity enables the IOSTAR™ toconduct missions which are not feasible using conventional satellites.These missions include, but are not limited to, satellite inspection,monitoring, rescue, retrieval, repair, servicing, refueling andrepositioning; direct communication services; in-orbit power generationfor other spacecraft like the International Space Station; andinterplanetary tasks, operations or transfers that may occur well beyondEarth orbit, such as trips to the Moon, the Asteroids, or the Planets.

The reactor 19 is also coupled to an energy converter 22 which convertsheat to electrical energy. In one embodiment, the energy converter 22includes a turbine driven by fluid that is heated by the reactor 19 toproduce a large amount of electrical power. The converter 22 is coupledto the boom 11, next to the radiation shield 20. An energy converter maybe an direct converter, which converts heat directly to electricity. Asan alternative, an energy converter may be an indirect converter, whichconverts thermal energy to mechanical energy, and then to electricalenergy. In a preferred embodiment of the invention, the converteremploys the Brayton Cycle. In alternative embodiments, the converter maybe a Rankine or Stirling Cycle converter. A thermoelectric or thermionicconverter may also be employed. In a preferred embodiment of theinvention, a recuperator may be connected to the energy converter.

II. Details of IOSTAR™ Embodiments

FIG. 3 provides a side view, where the IOSTAR™ is viewed along its sidein the plane of the radiator panels 16. FIG. 4 offers a view of theinvention in its fully collapsed configuration, capable of being stowedin a launch vehicle shroud 24.

FIGS. 5, 6, 7 and 8 present more detailed end and cross-sectional viewsof the IOSTAR™ stowed in the launch vehicle shroud 24. In one preferredembodiment of the invention, the IOSTAR™ is placed in orbit using theUnited States Space Shuttle.

FIG. 9 supplies a schematic block diagram of control systems 28 designedfor a preferred embodiment of the invention. A doubly redundant set ofCPUs manage the many subsystems aboard the IOSTAR™, including antennas30, docking and star cameras 32, 34, RADAR and LIDAR systems 36 fortracking objects or satellites 15, an ion thruster controller 38, andpower and thrust system controls 40. These systems enable the presentinvention to rendezvous and dock with a satellite or object in orbit. Inan alternative embodiment, the various sensors and cameras aboard theIOSTAR™ may be used to conduct remote sensing missions. The blockdiagram also relates the CPUs to attitude sensors and controls 42, the28 VDC power system 44, the bus health and attitude control subsystems46, 48 and an emergency blow down thruster control 50.

FIG. 10 offers a detailed schematic view of the ion propulsion system12. A mixture of helium and xenon flows from tank 13 to the ion engine12, where ions are created by a hollow cathode and accelerated through aseries of grids to provide thrust for the IOSTAR™ spacecraft.

FIG. 11 reveals a cross-sectional view of one embodiment of theinvention, depicting the launch vehicle shroud 24, radiators 16,manifolds 17 and energy converter 22.

FIG. 12 supplies a schematic diagram which offers an overview of theBrayton System, the energy converter 22 that is utilized in a preferredembodiment of the invention. Heat from the reactor 19 drives a turbine,which, in turn, drives an alternator and a compressor. A recuperatorincreases the efficiency of the system by recovering a portion of theheat from the turbine exhaust to pre-heat the working fluid. Radiators16 expel waste heat to outer space.

FIG. 13 provides a view of an alternative embodiment of the IOSTAR™which includes radiators disposed along the boom.

III. IOSTAR™ Missions & Operations

The present invention is different from conventional orbital systems, inthat it will be capable of accomplishing many missions over a long life.Although the IOSTAR™ will be reusable, in one embodiment the entiresystem will be capable of being launched using a single launch vehicle,preferably the United States Space Shuttle. Other launch vehicles thatare reusable or expendable may also be employed. The firstimplementation of the IOSTAR™ will be constructed primarily or entirelyon the Earth's surface, and then will be launched into orbit. Laterimplementations may be partially or completely constructed in orbit. Ingeneral, the IOSTAR™ may be controlled from a terrestrial operationscenter, or may operated by an on-orbit controller.

In general, the invention is fully extended after launch, and is thenready for operations. A first, general mission will comprise locating asatellite already in orbit, and then grasping, moving and releasing thatsatellite. IOSTAR™ will be able to move spacecraft between low Earthorbits and positions in higher orbits or to other locations in our SolarSystem. This primary mission of moving an object in space includestransporting satellites from one position in an orbit to another, fromone orbit to another, to distant locations beyond Earth orbit or fromdistant locations beyond Earth orbit back to Earth orbit. The IOSTAR™may be used for missions to the Moon, to the Planets or to theasteroids. Another mission may include changing the position of asatellite so that it is purposefully de-orbited.

In general, the term “rendezvous” pertains to the approach of an IOSTAR™to another object or objects in space. Rendezvous may or may includestation-keeping, or any contact, probing, interaction, coupling,observing or docking between an IOSTAR™ and another object.

Once the IOSTAR™ completes its rendezvous and docking with a satellite,the satellite may be transported for retrieval and/or repair. Ingeneral, the repositioning of a satellite from one location to anotherwill involve moving the satellite along an incremental, expanding,generally spiral pathway. FIG. 14 illustrates one of the basic methodsof the invention. A satellite 15 is first launched using a conventionalbooster to a low Earth orbit of roughly 150 nautical miles. The IOSTAR™10 then completes a rendezvous with the satellite 15, and engages thesatellite 15 with its docking hardware 14. The IOSTAR™ then graduallyraises the altitude of the satellite 15 to an operational orbit bymoving the payload along an incremental, expanding spiral pathway. Thisprocedure provides substantial cost savings for delivering a spacecraftto an operational orbit compared to the conventional technique oflaunching spacecraft with a multi-stage rocket. In an alternativeembodiment of the invention, the IOSTAR™ will be able to rendezvous withan object beyond Earth orbit. In this embodiment, the IOSTAR™ will becapable of retrieving an object or spacecraft from a remote locationbeyond Earth orbit.

FIG. 15 depicts an orbital repositioning mission. The invention may notonly be used to transport a new satellite to its destination orbit, butmay also be employed to capture a satellite which has reached the end ofits useful life and needs to be safely de-orbited or placed in adisposal orbit.

In general, the primary IOSTAR™ mission will involve rendezvousing anddocking with a spacecraft which is already in a low Earth orbit. Afterdocking, the IOSTAR™ will then move from a low Earth orbit to a highEarth orbit or to a position beyond Earth orbit. As an alternative, theIOSTAR™ will first travel to a high orbit or to a position beyond Earthorbit, locate and grasp an object, and then relocate it to Earth orbitor to a different position beyond Earth orbit.

FIGS. 16 & 17 compare a conventional geosynchronous mission to anIOSTAR™ mission. In a conventional launch, a satellite reaches highorbit in seven to ten hours, but at great expense. Using IOSTAR™, thesatellite takes a gradual spiral path over a 45 to 65 day period toreach high orbit, but at a much lower cost.

FIGS. 18, 19, 20 and 21 furnish generalized views of four representativeIOSTAR™ missions, including in-orbit placement, in-orbit repair,recovery and retrieval and Space Station Servicing. While all theIOSTAR™ objectives and missions are too numerous to delineate in thisSpecification, Table Three provides a representative and illustrativelist of uses for the present invention in outline form.

TABLE THREE Objectives & Missions Correct an anomalous satellite Earthorbit Provide mobility for a satellite in orbit Move a satellite inspace from one geostationary orbital position to another Move asatellite in space from one geosynchronous orbital position to anotherInspect a satellite in orbit Repair a satellite in orbit Extend usefullife of a satellite  By replenishing a consumable  By replenishing power By replenishing fuel  By replacing a battery  By replacing a satellitecomponent Reposition a satellite to a lower orbit Reposition a satelliteto a higher orbit Service a spacecraft in combination with the U.S.Space Shuttle Service a spacecraft in combination with the InternationalSpace Station Reposition a spacecraft from a low to a high orbit torealize cost savings compared to the costs of a  conventional launchMove a satellite into a disposal orbit Provide services to an insurerSalvage a spacecraft in accordance with an insurance contract Enable aninsurer to lower launch premiums Obtain information about a failure ofan orbiting asset or spacecraft Enable an insurer to lower the financialrisks of a spacecraft launch Maintain a fleet of operating spacecraft,including United States Global Positioning Satellites Supply on-orbitpower to another spacecraft Move spare spacecraft from one orbitalaltitude or plane to another Provide services to a spacecraftmanufacturer Provide services to a spacecraft user Provide services to agovernment agency Use IOSTAR ™ as a reusable upper stage of aconventional launch vehicle to reduce launch costs Use IOSTAR ™ and alaser used for orbital debris removal Use laser to divert an asteroidProduce propellant from an asteroid Produce propellant from waterlaunched into orbit from Earth Produce propellant from a stable,storable material launched into orbit from Earth Process ice present onan asteroid by electrolysis to form hydrogen and oxygen Processcarbonaceous material present on an asteroid to form a storablepropellant Recycle satellites in space

FIGS. 22 and 23 portray the IOSTAR™ in combination with theInternational Space Station. One embodiment of the invention will beconfigured to provide direct communication services that include any oneor two-way transmissions or emanations between or among the IOSTAR™ andterminals on or near the Earth's surface, or with other satellites orspacecraft. One example of a conventional direct communication serviceis a high-bandwidth transmission to consumers like DirecTV™. In general,these direct communication services will be conducted usingelectromagnetic, optical or any other suitable frequencies or modes ofcommunication over a distance. In one embodiment of the invention,IOSTAR's direct communication services will be conducted using frequencybands 11and 12. Frequency band 11 extends from 30 to 300 GigaHertz, andis also referred to by the term “millimetric waves.” Frequency band 12extends from 300 to 3000 GigaHertz or 3 TeraHertz, and is also referredto by the term decimillimetric waves. This nomenclature of frequencybands was adopted in the Radio Regulations of the InternationalTelecommunication Union, Article 2, Section 11, Geneva, 1959.

These direct communication services will generally be enabled byIOSTAR's enormous power generating capabilities. FIG. 24 is aperspective view of a satellite having an array of antennas which may beused in combination with an IOSTAR™ to provide a direct broadcast system52. Since the IOSTAR™ can generate very high levels of electrical powercompared to conventional satellites 15, it may be used to transmitdirect broadcast signals at extremely high frequencies. The Ka-Band(20-30 GHz) is the highest range of radio frequencies that are currentlyused by commercial satellites to communicate with customers on theground. By drawing on its massive power supply, the IOSTAR™ DirectBroadcast System will be capable of offering regulated direct broadcastsignals at frequencies of 100 GHz and beyond using high poweramplifiers, such as a traveling wave tube amplifiers or grid amplifiers.This direct broadcast system may also employ and a beam-forming array ora steerable antenna to penetrate layers of the atmosphere which absorband scatter conventional, lower power signals. In general, the presentinvention is capable of generating a vast amount of electrical power toprovide a wide variety of direct communication services that offerdirect transmissions between the present invention and terrestrialterminals. In one embodiment of the invention, direct communicationservices are conducted using frequency bands 11 and 12. In general,these direct communication services may be provided by the presentinvention utilizing any means, mechanism or phenomenon that exploitsparticle or electromagnetic wave transmissions, forces, fields or actionat a distance, including the radio-frequency and optical spectra.

IV. Business Methods

A. Services

The present invention furnishes three general categories of services:

-   -   Providing controlled kinetic energy;    -   Generating and/or providing power; and    -   Sensing and/or gathering and/or supplying information or data.        These services are depicted generally in FIG. 25.        Kinetic Energy

In this Specification and in the Claims that follow, the term“controlled kinetic energy” relates to the conversion of a fuel tomechanical energy on board a vehicle in orbit to accomplish specifictasks. The control of the kinetic energy may originate from a groundstation on the Earth, from another satellite, or from another celestialbody. In one embodiment, the conversion of fuel involves fission in anon board nuclear reactor. The term “operating a nuclear powered vehiclein orbit” connotes a series of extended and perhaps varied missions overa relatively long vehicle lifetime which is enabled by the nuclearreactor. Examples of services pertaining to controlled kinetic energyinclude, but are not limited to:

-   -   Supplying controlled kinetic energy to another satellite, which        may be used to move, rescue, repair, resupply, salvage or        otherwise interact or impinge upon a satellite    -   Supplying controlled kinetic energy to transport a payload to or        from a satellite or celestial body

When used in this Specification and in the Claims that follow, thedelivery of a payload “to” or “from” a celestial body is intended toinclude delivery to an orbit around that celestial body, or delivery toor near the surface of that celestial body.

Power

In this Specification and in the Claims that follow, the term“electrical energy” relates to the generation and/or storage and/or useof a stored quantity, current or flow of electricity derived fromelectromagnetic forces. Examples of services pertaining to electricalenergy include, but are not limited to:

-   -   Using electricity generated by an IOSTAR™ to move, rescue,        repair, resupply, salvage, interact or otherwise impinge upon a        satellite    -   Using electricity generated by an IOSTAR™ to transport a payload        to or from a satellite or celestial body

The electrical energy generated on board the IOSTAR™ is generated by aturbine or some other heat conversion device. The heat is produced bythe nuclear reactor.

The electrical energy from IOSTAR™ may be conveyed to another satelliteusing a direct connecting link like a cable, or may be beamed to anothersatellite or to a celestial body, including Earth, in the form of aradiated energy beam, wave or stream of particles. In an alternativeembodiment, electricity may be stored aboard the IOSTAR™ in a storagedevice, such as a battery or a fuel cell. The battery or fuel cell maythen be physically delivered to another satellite or to a celestialbody. The electrical energy may be provided to one or to a plurality ofcustomers. A number of customers may be served on a time-share basis.

Information or Data

In this Specification and in the Claims that follow, the terms“information” and “data” relate to any manifestation, form or type ofintelligence, pattern, language, image, content, audio, video,communication, or any other type of sensation. Examples of servicespertaining to information or data include, but are not limited to:

-   -   Processing information    -   Sensing information    -   Relaying information    -   Measuring radiation    -   Imaging    -   Environmental studies    -   Natural resource studies    -   Border surveillance    -   Reconnaissance    -   Conveying information to another satellite or celestial body,        including Earth    -   Direct broadcast services    -   Two-way telecommunications    -   Using sensed or received information to move, rescue, repair,        resupply, salvage, interact with or otherwise impinge upon a        satellite

In this Specification and in the Claims that follow the term “processinginformation” is used to connote any kind or type of information or dataprocessing, including, but not limited to, sensing, conveying, relaying,distribution, switching, transmitting, receiving, storing or computing.Sensed or received information may be necessary to locate a satellite,rendezvous and then provide the required service. Sensed or receivedinformation may also be used for navigation purposes to transport apayload to another satellite or celestial body. Information services maybe provided to a single customer, or to a plurality of customers.

B. Transactions

The present invention may be involved in three general categories ofbusiness transactions:

A Sale

A Contract for Limited Use or Lease

A Contract for Payment by Unit or Rate

Each of these transactions may involve the services described above,including providing controlled kinetic energy; generating and/orproviding power; and gathering and/or supplying information or data.

A sale or purchase is generally a transfer of property in which fulltitle passes to the customer. Another alternative transaction is a tradefor some other goods and/or services. In accordance with this categoryof transaction, a government agency or private company might purchase anIOSTAR™.

A Contract for Limited Use is generally a license or permission to useor utilize for a predetermined purpose or time, in which less than fulltitle is conveyed to a customer. As an example, a Contract for LimitedUse may be authorize a lease of an IOSTAR™ for the time it takes tocomplete a particular objective or task, such as a satellite deploymentor rescue, or for a defined time period, such as a month or a year.

A Contract for Payment by Unit or Rate generally requires a customer tomake a payment based on agreed upon terms involving a quantity of goodsor services. As examples, this type of transaction might call for apayment based on:

-   -   An amount of electrical power supplied from an IOSTAR™ to        another satellite (in Watt-Hours or Amp-Hours).    -   A quantity of mass transported by an IOSTAR™ to the Space        Station or to the Moon (measured in kg, or measured by the        distance moved in km).    -   A change in a satellite's orbital parameter, such as, but not        limited to, altitude, apogee, perigee or inclination.    -   A number of packets of data or information conveyed to a        customer or some other recipient, such as programs of content        conveyed to receivers on Earth by an IOSTAR™ direct broadcast        service.    -   An illuminated power flux density originating from an IOSTAR™.

These services may be provided to a single customer, or generallysimultaneously to a plurality of customers. Services may also beprovided in a time-share arrangement.

C. Service Provided to Satellites and Celestial Bodies

In general, all services and transactions concerning the presentinvention involve another satellite or a person, receiver, terminal,building, vehicle or other object on or near a celestial body, or thecelestial body itself.

V. Propagated Signals

Another class of embodiments of the invention comprise propagatedsignals. The IOSTAR™ may be used as the source of power or informationsignals that are radiated to other satellites or to celestial bodies.

Scope of the Claims

Although the present invention has been described in detail withreference to a particular, preferred embodiment and alternateembodiments, persons possessing ordinary skill in the art to which thisinvention pertains will appreciate that various modifications andenhancements may be made without departing from the spirit and scope ofthe Claims that follow. The various embodiments, implementations andapplications that have been disclosed above are intended to educate thereader about particular embodiments, and are not intended to constrainthe limits of the invention or the scope of the Claims. The List ofReference Characters which follows is intended to provide the readerwith a convenient means of identifying elements of the invention in theSpecification and Drawings. This list is not intended to delineate ornarrow the scope of the Claims.

LIST OF REFERENCE CHARACTERS 10 In-Orbit Space Transportation & RescueSystem, or IOSTAR ™ 11 Collapsible spacecraft boom 12 Electricpropulsion system 13 Propellant tank 14 Grasping/Docking mechanism 15Satellite or other payload or cargo 16 Radiator 17 Manifold bellows 18Gas flow tubes 19 Nuclear reactor 20 Radiation shield 22 Energyconverter 24 Launch vehicle 28 Block diagram of control systems 30Antenna 32 Docking cameras 34 Star cameras 36 RADAR & LIDAR 38 Ionthruster controller 40 Power and thrust system controllers 42 Attitudesensors and controls 44 28 VDC charger and regulator 46 Bus health andstatus multiplexer and D/A converters 48 Attitude control thrusteron/off control 50 Emergency blowdown thruster control 52 IOSTAR ™ DirectBroadcast System

1. A method comprising the steps of: a) operating a nuclear poweredvehicle in outer space wherein the nuclear powered vehicle is providedwith a boom, docking hardware proximate a first end of said boom used tocapture a satellite, a nuclear reactor coupled to said boom proximate asecond end of said boom opposite the first end of said boom, a radiationshield coupled to said nuclear reactor, and an energy converter coupledto the nuclear reactor, the nuclear reactor and energy converterproducing at least 250 kW of sustained electrical power; and b)processing information on board said nuclear powered vehicle to rescue asaid satellite, wherein the nuclear powered vehicle travels to andcaptures the satellite from a first orbit and transfers the satellite toa second orbit, and wherein the nuclear powered vehicle approaches,connects to and transits with the satellite such that the radiationshield protects the satellite from radiation from the nuclear reactor.2. A method of operating a nuclear powered vehicle in orbit comprisingthe steps of: a) providing said nuclear powered vehicle in orbit with aradiation shield for protecting a detachable payload; b) providing saidnuclear powered vehicle with a grasping means extending outwardlytherefrom at one end for docking and interacting with at least one othersatellite; c) said grasping means being configured for multiple use andfor interacting with a plurality of different objects; d) said graspingmeans including a plurality of segments; e) configuring said pluralityof segments to partially surround one of said plurality of differentobjects and to engage and to grasp said object without the need for anypreconfigured docking interface on said object; and f) processinginformation; said processed information for interacting with at leastone other satellite.
 3. A method as recited in claim 2, in which saidinformation is used to affect another satellite.
 4. A method as recitedin claim 2, in which said information is used to move a satellite.
 5. Amethod as recited in claim 2, in which said information is used totransport a payload to a satellite.
 6. A method as recited in claim 2,in which said information is used to transport a payload from asatellite.
 7. A method as recited in claim 2, in which said informationis used to transport a payload to a celestial body.
 8. A method asrecited in claim 2, in which said information is used to transport apayload from a celestial body.
 9. A method as recited in claim 2, inwhich said nuclear powered vehicle for providing said information issold.
 10. A method as recited in claim 2, in which said nuclear poweredvehicle for providing said information is leased for a specified task.11. A method as recited in claim 2, in which said nuclear poweredvehicle for providing said information is leased for a specified time.12. A method as recited in claim 2, in which a customer who uses saidnuclear powered vehicle for providing said information is chargedaccording to a specified rate.
 13. A method as recited in claim 2, inwhich said information is conveyed to another satellite.
 14. A method asrecited in claim 12, in which said information is conveyed to a receivergenerally near a celestial body.
 15. A method as recited in claim 2, inwhich said information is conveyed using a radio signal.
 16. A method asrecited in claim 2, in which said information is provided to a pluralityof customers.
 17. A method as recited in claim 2, in which a customer ischarged for receiving said information by the number of packetsconveyed.
 18. A method as recited in claim 2, in which said informationis used for reconnaissance.
 19. A method as recited in claim 2, in whichsaid information is used for surveillance.
 20. A method comprising thesteps of: a) operating a nuclear powered vehicle in outer space whereinthe nuclear powered vehicle is provided with a boom, docking hardwareproximate a first end of said boom used to capture a satellite, anuclear reactor coupled to said boom proximate a second end of said boomopposite the first end of said boom, a radiation shield coupled to saidnuclear reactor, and an energy converter coupled to the nuclear reactor,the nuclear reactor and energy converter producing at least 250 kW ofsustained electrical power; and b) processing information on board saidnuclear powered vehicle to repair said satellite, wherein the nuclearpowered vehicle travels to and captures the satellite from a first orbitand transfers the satellite to a second orbit where the satellite isrepaired, and wherein the nuclear powered vehicle approaches, connectsto and transits with the satellite such that the radiation shieldprotects the satellite from radiation from the nuclear reactor.
 21. Amethod of rescuing a satellite comprising the steps of: a) operating anuclear powered vehicle in outer space to capture the satellite while itis in a first orbit, wherein the nuclear powered vehicle is providedwith a boom, docking hardware proximate a first end of said boom used tocapture the satellite, a nuclear reactor coupled to said boom proximatea second end of said boom opposite the first end of said boom, aradiation shield coupled to said nuclear reactor, and an energyconverter coupled to the nuclear reactor, the nuclear reactor and energyconverter producing at least 250 kW of sustained electrical power; andb) using the nuclear powered vehicle to move the satellite from itsfirst orbit to a second orbit, and wherein the nuclear powered vehicleapproaches, connects to and transits with the satellite such that theradiation shield protects the satellite from radiation from the nuclearreactor.
 22. The method of claim 21 wherein the nuclear powered vehicleis further provided with an ion propulsion system coupled to said boom,a propellant tank storing fuel for the ion propulsion system coupled tosaid boom, and a radiator, and the energy converter comprises a turbine,a compressor and a generator.
 23. The method of claim 21 wherein thefirst orbit is a geosynchronous orbit.
 24. The method of claim 23wherein the second orbit is a low Earth orbit.
 25. A method comprisingthe steps of: a) operating a nuclear powered vehicle in outer space tocapture a satellite while it is in a first orbit, wherein the nuclearpowered vehicle is provided with a boom; docking hardware proximate afirst end of said boom used to capture the satellite; a nuclear reactorcoupled to said boom proximate a second end of said boom opposite thefirst end of said boom; a radiation shield coupled to said nuclearreactor; and an energy converter coupled to the nuclear reactor, thenuclear reactor and energy converter producing at least 250 kW ofsustained electrical power; and b) using the nuclear powered vehicle tomove the satellite from its first orbit to a second orbit, and whereinthe nuclear powered vehicle approaches, connects to and transits withthe satellite such that the radiation shield protects the satellite fromradiation from the nuclear reactor.